As Federal noise level standards for gas turbine engines become more stringent, the designer of such engines is faced with the problem of reducing noise in an efficient, cost effective manner.
Gas turbine engine noise is generated from two primary sources. First, there is that associated with the viscous shearing of rapidly moving gases exhausted into the surrounding atmosphere. In turbofan aircraft engines, such gases are emitted from the fan and core nozzles at the rear of the engine. Various approaches have been utilized to reduce this shear noise, most approaches incorporating mixers to co-mingle fan and exhaust gases with each other and with the surrounding environment.
The second source of noise, and the one to which the present invention is directed, is generated by the rotating turbomachinery itself. This results from the relative motion between the rapidly rotating blade rows and the interflowing gas stream. The noise is affected by such parameters as blade rotational speed, blade-to-blade spacing, blade geometry, and by the proximity of stationary hardware to such rotating blade rows, as in the case of an outlet guide vane arrangement. Another example of the latter condition occurs in typical multistage axial compressors where stationary blade rows are alternated with rotating blade rows. Some of the noise generated in this manner can be absorbed and suppressed by means of acoustic or sound-absorbing paneling disposed about the periphery of the nacelle enclosing the rotating turbomachinery. Such sound absorbing material is well known in the art. However, because of the close proximity of the fan or compressor to the inlet frontal plane, and the lack of acoustic shielding in the forward direction, a significant percentage of noise propagates forward from the gas turbine inlet duct.
Prior attempts to solve this problem have concentrated on the application of sound-absorbing material to the inlet duct inner walls. This does little to attenuate unreflected noise propagating in the axially forward direction. Additional benefits have been obtained by providing coaxial, circumferential rings of sound-absorbant material within the inlet. However, such rings produce a loss of inlet total pressure and, therefore, bring about performance losses which remain throughout the engine operating envelope even when noise propagation presents no hazard or nuisance to inhabitants below. Another disadvantage is that anti-icing provisions must be made. Such a structure increases the potential for foreign object damage, decreases accessibility to the fan rotor on the flight line and increases weight.
Another concept incorporates an axially translating noise deflector on the bottom of the inlet duct to selectively reduce the downward transmission of noise from the inlet. However, this configuration is inadequate for two reasons. First, it has been shown that an inlet incorporating such a deflector may have a poor pressure recovery characteristic (i.e. it is inherently a high-loss system) depending upon its shape. Secondly, and somewhat related to the foregoing problems, is that the total pressure pattern may become highly distorted, as for example in the plane of a gas turbine fan stage disposed within the duct. While the poor pressure recovery results in degraded engine performance, the distorted flow pattern may, under certain conditions, cause excessive fan blade stresses and possible destruction of the rotating turbomachinery.
Yet another approach which has recently been investigated is to accelerate the inlet flow such that the average Mach number at the throat is 1. The principle employed is that an acoustic wave cannot propagate upstream against a Mach 1 flow since the wave itself can only travel at Mach 1. This, however, presents some performance problems. For application of this concept to conventional aircraft, a considerable amount of inlet area variation is required because of the large variation in air flow with engine power setting. Further, the inlet length must be increased considerably due to the necessity of diffusing from Mach 1 to an acceptable fan rotor plane Mach number without boundary layer separation in the inlet. Additionally, recent tests have shown that there is a serious loss of noise suppression when a choked inlet is subjected to inflow at various angles of attack within the normal take-off and landing range.
Prior attempts to combine the acoustic benefits of a choked, or near-choked, inlet with those of the sound-absorbing material have been disappointing in that the system was no more effective than that of the choked inlet alone.
The problem facing the gas turbine designer is, therefore, to provide a means for attenuating noise emanating from the duct without incurring excessive complexity, cost or performance penalties.